The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor to generate hot combustion gases. Energy is extracted from the gases in a turbine which powers the compressor and produces useful work.
The turbine includes a row of rotor blades extending outwardly from a supporting disk, with each blade having an airfoil configured for extracting energy from the gases to rotate the disk. The airfoil has pressure and suction sides extending between leading and trailing edges and from root to tip. The airfoil tip is spaced radially inwardly from a stationary shroud to define a small gap therebetween. The gap is sized as small as practical to minimize the amount of combustion gas leakage therethrough for maximizing engine efficiency. However, differential expansion and contraction between the rotor blades and the stationary shroud occasionally permit tip rubs which must be accommodated.
Since the blade airfoil is bathed in hot combustion gases during operation, it is typically cooled by channeling therethrough a portion of air bled from the compressor. The airfoil is hollow and includes one or more cooling circuits therein which can have various configurations, and pins and turbulators for enhancing heat transfer of the cooling air therein. The airfoil typically includes rows of discharge holes through the sidewalls which produce cooling air films on the external surface of the airfoil for protection against the hot combustion gases.
However, the airfoil tip is particularly difficult to effectively cool since it is closely spaced near the shroud and is subject to combustion gas flow therebetween and occasional tip rubs.
Accordingly, a typical turbine blade tip includes a squealer tip rib which extends around the perimeter of the airfoil flush with its sides and defines a tip cavity and a floor therebetween. The tip rib reduces the surface area between the tip and shroud subject to tip rubbing, but is subject to heating from the three exposed sides thereof. Cooling air may be discharged through an axial row of film cooling holes below the pressure side tip rib for cooling thereof, and additional discharge holes may be provided through the tip floor for discharge into the tip cavity.
Since the airfoil tip varies in thickness between the leading and trailing edges, the effectiveness of the pressure side film cooling air is limited. As the film cooling air travels over the pressure side tip rib, it encounters combustion gas leaking through the tip gap. Recirculation of the cooling air and combustion gas within the tip cavity reduces the cooling effectiveness of the air in the tip gap.
Accordingly, it is desired to provide an improved turbine airfoil tip configuration having enhanced cooling for improving blade life.